Electric sail for producing spacecraft propulsion

ABSTRACT

A spacecraft propulsion system includes a plurality of wires ( 102 ) or other electrically conductive elongated members deployed from a main body ( 101 ) into respective radial directions. An electric potential generator ( 605 ) generates an electric potential on board the main body ( 101 ). The electric coupling between the electric potential generator ( 605 ) and the elongated members is controlled ( 604 ) so that all or some of the elongated members ( 102 ) assume a high positive potential. An auxiliary propulsion system ( 203 ) rotates the main body around a rotational axis ( 502 ) that is perpendicular to the radial directions, thus creating a centrifugal supporting force to the elongated members.

TECHNICAL FIELD

The invention concerns generally the technology of spacecraftpropulsion. Especially the invention concerns the technology ofutilizing solar wind as the source of propulsive force.

BACKGROUND OF THE INVENTION

A spacecraft propulsion system is a subsystem of a spacecraft whosepurpose is to change the state of motion of the spacecraft from itsnatural Keplerian motion. The Keplerian motion is due to the gravityfield of the solar system bodies. Among the figures of merit of apropulsion system are the payload mass fraction and the delta-v thepropulsion system can produce. Larger is better for both figures ofmerit. The payload mass fraction is the payload mass divided by thetotal initial mass (payload mass plus initial propulsion system mass) ofthe spacecraft. The delta-v is the time integral, computed over theworking time of the propulsion system, of the non-gravitationalacceleration provided by the propulsion system.

Conventional propulsion systems include chemical rockets and electricpropulsion. For payload mass fraction of ⅓ the best chemical rocketbipropellant (liquid hydrogen plus liquid oxygen) provides a delta-v ofabout 1 AU/year, where AU means one astronomical unit, essentially equalto 150 million kilometres. Higher delta-v values are possible but withexponentially decreasing payload mass fraction. For electric propulsionsystems a firmer delta-v value does not exist, but typical values forrealized missions are 2-4 AU/year. These delta-v values are notsufficiently high for many purposes, e.g. for reaching outer solarsystem targets in a reasonable time. While somewhat higher delta-vvalues can be generated by reducing the payload mass fraction to aminimum, for a fixed payload this means an exponential growth of theinitial mass and corresponding exponential increase of the mission cost.

For producing propulsive force alternative solutions exist that takeadvantage of naturally occurring phenomena in space. A solar sail is alarge sheet of thin membrane that the spacecraft deploys once outsidethe Earth's atmosphere. Photons originating from the Sun hit the sail asa continuous stream, thus transferring momentum to it. A magnetic sailconsists of one or more large-area loops of (preferably superconductive) wire, through which an electric current is driven in orderto create a magnetic field. The field interacts dynamically with thesolar wind and generates a propulsive force. A magnetic sail is knownfrom the prior art publication R. M. Zubrin, D. G. Andrews: “Magneticsails and interplanetary travel”, Journal of Spacecraft and Rockets,Vol. 28, No. 2, pp. 197-203, published in 1991.

The idea of using solar wind for generating propulsion is also knownfrom the publication P. Janhunen: “Electric Sail for SpacecraftPropulsion”, Journal of Propulsion and Power, Vol. 20, No. 4, pp.763-764, published in 2004 and incorporated herein by reference. Solarwind means the continuous stream of charged particles, mostlyhigh-energy electrons and protons, that the Sun emits continuously toessentially all radial directions. An electric sail is an electricallyconductive structure that is held at a positive potential with respectto the solar wind plasma. In contrast to the solar sail which extractsmomentum from the Sun's electromagnetic radiation and not the solarwind, the electric sail does not need to be a continuous sheet. Saidprior art publication presents an example in which the electric sail isa mesh of wires with a spacing less than or equal to the so-called Debyelength of the plasma. The Debye length is a measure of the distance overwhich an Individual charged particle can exert an effect.

Although the usability of an electric sail for spacecraft propulsion hasthus been theoretically shown, there are no known practical solutionsthat would be applicable to implement the principle on actual spacemissions.

SUMMARY OF THE INVENTION

An objective of the invention is to present a practically applicableelectric sail system. Another objective of the invention is to present aspacecraft propulsion subsystem that facilitates easy and reliabledeployment of the electric sail. A further objective of the invention isto present a method and a system with which it is possible to steer anelectric-sail-propelled spacecraft. Yet another objective of theinvention is to present an electric sail system that is tolerant againstmicrometeors and other hazards caused by the space environment.

The objectives of the invention are achieved by making the electric sailcomprise a number of radially extending multifilament wires that areheld tight by the centrifugal force.

A spacecraft propulsion system according to the invention comprises:

-   -   a plurality of electrically conductive elongated members adapted        to be deployed from a main body Into respective radial        directions,    -   an electric potential generator adapted to generate an electric        potential on board said main body,    -   a controllable electric coupling between said electric potential        generator and said plurality of electrically conductive        elongated members, and    -   an auxiliary propulsion system adapted to rotate said main body        around a rotational axis that is perpendicular to said radial        directions.

The Invention applies also to a method for changing the state of motionof a spacecraft from its natural Keplehan motion. The method accordingto the invention comprises:

-   -   generating an electric potential on board a main body of the        spacecraft.    -   controlling an electric coupling between the generated electric        potential and a plurality of electrically conductive elongated        members deployed from said main body into respective radial        directions, and    -   rotating said main body around a rotational axis that is        perpendicular to said radial directions in order to cause a        centrifugal tensile force to said plurality of electrically        conductive elongated members.

Additionally the invention applies to a computer program product on acomputer-readable medium. A computer program product according to theinvention comprises software instructions that, when executed In acomputer System, cause the implementation of:

-   -   controlling the generation of an electric potential on board a        main body of a spacecraft,    -   controlling an electric coupling between the generated electric        potential and a plurality of electrically conductive elongated        members deployed from said main body into respective radial        directions, and    -   controlling the rotation of said main body around a rotational        axis that is perpendicular to said radial directions in order to        cause a centrifugal tensile force to said plurality of        electrically conductive elongated members.

In the invention we utilise the fact that an electric sail does not needto be a continuous membrane like a solar sail, but even a relativelysparse mesh will do. Additionally we utilise the principle of theso-called centrifugal force, which means that a mass fixed to a rotatingobject and not located on the axis of rotation will create a constantpulling force in the radial direction. Deploying -a number of wiresextending radially into space from a spinning spacecraft will result ina spinning, cartwheel-like configuration, where the centrifugal forcekeeps the radially extending wires tight.

If the wires are electrically conductive, it is possible to keep them ata positive potential relative to the surrounding plasma by coupling themelectrically to an electron gun on board the spacecraft. The positivepotential of the wires causes an electrostatic Coulomb force interactionwith the protons of the solar wind, which in turn transfers momentumfrom the solar wind to the wires and therethrough to the wholespacecraft. Drawing electrons from the wires to the spacecraft body andemitting them to space with the electron gun counteracts the effects ofsolar wind electrons, which would otherwise neutralise the positivepotential. Many kinds of steering manoeuvrea become possible, if therate at which electrons are drawn from each wire is controllable.

Advantageous embodiments of the invention are discussed in the dependingclaims. The features recited in depending claims are mutually freelycombinable unless otherwise explicitly stated. The exemplary embodimentsof the invention presented in this patent application are not to beinterpreted to pose limitations to the applicability of the appendedclaims. The verb “to comprise” is used in this patent application as anopen limitation that does not exclude the existence of also unrecitedfeatures.

The novel features which are considered as characteristic of theinvention are set forth in particular in the appended claims. Theinvention itself however, both as to its construction and Its method ofoperation, together with additional objects and advantages thereof, willbe best understood from the following description of specificembodiments when read in connection with the accompanying drawings. Thedrawings are schematic only and not drawn to scale.

BRIEF DESCRIPTION OF DRAWINGS

FIG. 1 illustrates the principle of a centrifugally supported electricsail,

FIGS. 2 a and 2 b illustrate certain aspects of wire deployment,

FIG. 3 illustrates the principle of solar wind assisted spinning,

FIG. 4 illustrates certain structural considerations of attaching thewires,

FIGS. 5 a, 5 b and 5 c illustrate manoeuvres with an electric sail, and

FIG. 6 illustrates an exemplary functional diagram of certain parts of aspacecraft with electric sail propulsion.

DETAILED DESCRIPTION OF THE INVENTION

Principal Configuration

FIG. 1 illustrates schematically the basic principle of an electric sailthat consists purely of radially extending wires, with no perimeterconnection or other interconnections between the wires. In the middlethere is the main body 101 of the spacecraft. Extending radiallyoutwards from the main body 101, there are a number of wires, of whichwire 102 is an example. The whole system rotates (spins) around arotational axis that is perpendicular to the plane of the paper and goesthrough the contra of gravity of the system, which we assume to coincidewith the center of the main body 101. The system is in a spaceenvironment, which means that there is essentially no similar dragcaused by any surrounding medium as for example in the Earth'satmosphere. On the other hand, since each noncentrally located piece ofa system rotating at constant angular velocity requires a constantcentripetal force to keep It on its circular orbit, the result of thespinning movement is that all wires are kept tight and point essentiallydirectly to the respective radial direction.

We also assume that the main body 101 of the spacecraft comprises anelectron gun or corresponding means for controllably emitting a stream103 of negative charges (essentially: electrons) to space. The stream ofnegative charges can basically be emitted to any direction, for examplealong the spin axis of the spacecraft to the downstream direction of thesurrounding plasma flow. Emitting negative charges will cause the mainbody 101 to assume a positive potential with respect to the surroundingplasma. If the wires are electrically conductive and have anelectrically conductive connection to the main body 101, electrons flowfrom the wires to the main body as is shown on wire 104 and the wirestoo assume a positive potential. The overall electric field of theSpacecraft shown in FIG. 1 starts to resemble that of a positivelycharged conductive disc. A minimum useful potential for the wires isabout 1 kV, which corresponds to the typical kinetic energy of solarwind protons. Higher values like 8-20 kV are likely to be morebeneficial.

The solar wind consists of protons and electrons moving essentiallyradially outwards from the Sun at high speed, typically 400-800 km/s.The disc-like electric field of the spacecraft repels the positivelycharged protons in accordance with Coulomb's law, which means that thecontinuous stream of solar wind protons pushes the spacecraft much likethe ordinary wind would push a circular sheet of windproof material inthe Earth's atmosphere. The electrons of the solar wind are attracted bythe positive potential of the wires (and of that of the main body 101),but since the momentum carried by electrons is negligible compared tothat carried by protons, essentially the only noticeable electron effectis the tendency of the gathered electron current to neutralise thepositive potential. The efficiency of the apparatus used to emit theelectron stream 103 from the spacecraft must be high enough tocounteract the neutralising effect of the gathered electron current.

Wire Material and Construction

Experiments and calculations suggest that a single thin monofilamentwire 100 meters long would only survive some months in space beforegetting severed by a micrometer. The “wires” that constitute theelectric sail must therefore not be monofilament wires but have aconstruction that has better chances to survive. A number of suitableconstruction principles have been suggested for other purposes in spacetechnology. Various kinds of braids, multifilament wires, cables, ropesand tethers can be used. One possible construction is the one introducedas Hoytether in a reference publication U.S. Pat. No. 6,286,788 B1. Thewires can also have the appearance of a band. For simplicity, only theterm “wire” is used throughout this description, but all other elongatedmember constructions are encompassed as well. The operating principle assuch would not exclude even rigid rod- or beam-like constructions, butit would be extremely difficult to construct an ultra-light-weightdeployable structure of the required dimensions from rigid pieces.

The relatively high positive potential and electric conductivity of thewire means that if and when it consists of a number of separatefilaments or component strings, these are all in the same potential andrepel each other. Thus the filaments or component strings are naturallykept apart, which reduces the risk that a micrometer would cut all ofthem simultaneously.

The wire material should have high tensile strength, low density andgood electrical conductivity. Good material choices include steel alloysand other metals that have high tensile strength. The wire material mayalso be a composite, like a carbon fiber or aramid fiber core with asurface metallization or other electrically conductive coating, or ametal core wire with fiber coating.

The wire should be as thin as possible, not only in order to save massand space in the spacecraft before electric sail deployment but also Inorder to keep the gathered electron current constituted by the solarwind elections as low as possible. The electron current is approximatelyproportional to the outer surface area of the wires.

How long the wires will be depends on many factors, like the totalnumber of wires, spacecraft mass, desired magnitude of propulsion,intended orbit radius (i.e. distance from the Sun) and the like. As arough assumption, the wires could be for example 10 kilometres long. Anultimate limit of the length of the wires comes from the tensilestrength of the wire material: each segment of wire must stand thecentripetal force (plus a safety marginal) that the remaining portion ofwire between that segment and the distant end will cause. It isnaturally possible to use wires that have non-constant cross-section, sothat the tensile strength of the wire would be a decreasing function ofthe distance from the main body of the spacecraft towards the distantend of the wire. All wires need not be of equal length.

Deployment Procedure

Initially the wires or wire groups are stored on reels, FIG. 2 a showsschematically an exemplary configuration, where the reels 201 arelocated at the edge of the generally disk-like main body 101 of thespacecraft. When the spacecraft Is set to rotation, the wires unwindthemselves assisted by the centrifugal force. Depending on the wireproperties and the reel diameter, the unwound wire may have a tendencyto curl up slightly. If this threatens to prohibit the windup processfrom starting properly, the use of a small ballast mass 202 at the endof the wire fixes the situation. The ballast mass 202 needs to be only afraction of the total mass of the ire, so its effect on the mass budgetof the propulsion system is minimal.

If the reels are unwound at constant speed, the wires are straightduring the deployment process, but inclined in an angle to the radialdirection. The angle depends on the rotation speed, unwinding speed andradial distance of the support point of the wire from the center ofrotation of the system; the dependency is of the form sinψ=2v/Rω, whereψ is the angle between the wire and the radial direction, v is the speedat which the wire is deployed, R is the radial distance of the supportpoint end ω is the angular speed of the system. The term “system” meanshere the whole mechanical system, i.e. the main body 101 of thespacecraft and the wires which are being unwound. FIG. 2 b illustratessaid quantities graphically.

During deployment, the system's total angular momentum tends to stayconstant, which means that the rotation rate of the system tends todecrease when the wire unreeling process proceeds. Thus, either thedeployment must begin when the spacecraft is spinning at a relativelyhigh rate, which decreases during the unreeling process and eventuallyreaches a suitable final value when the wires have been fully deployed,or a torque must be applied continuously or intermittently duringdeployment to increase the angular momentum. A straightforward way ofobtaining this torque is to use an auxiliary, conventional propulsionsystem. To maximise the torque, the value of which is thrust force timesarm length, the propulsive units 203 of the auxiliary system should beplaced as far as possible from the center of rotation of the system,e.g. at ends of long propulsion arms 204. The propulsive units 203 canbe e.g. small chemical rockets, pressure release thrusters, ion orplasma engines or any other suitable propulsion-generating means.

In the exemplary embodiment of FIG. 2 a the auxiliary propulsion systemcomprises a propulsion platform 205 and its attachment 206 to the mainbody 101 of the spacecraft. One possibility is to make the attachment206 rotatable with a controllable mechanical interaction with the mainbody. In such an embodiment the propulsive units 203 would be used tospin up the auxiliary propulsion system, which would rotate in relationto the main body 101, and use said controllable mechanical interactionto controllably administrate angular momentum from the spinningauxiliary propulsion system to the main body 101.

The auxiliary propulsion system and Its propulsion arms 204 are notneeded after the deployment phase is complete, so if wanted, they can bedetached (jettisoned) to increase performance during the propulsivestage. If flight-time wire length control is not used, the wire reels201 are also not needed after the wires have been unwound. Therefore,they can also belong to the detached system. In such a case, a mechanismmust exist for re-attaching the wires from the reels to the main body101 of the spacecraft. The eventual rotation rate of the spacecraft withfully deployed wires might be for example one whole turn in every 5minutes, but this is an exemplary value only and does not limit theapplicability of the invention. The rotation rate must be selected sothat taken the maximum expected rate of solar wind and the voltage valueof the wires, the centrifugal motion is still sufficient to keep thewires from bending by more than some suitable limiting value.

Besides by conventional propulsion, one may also add angular momentum tothe system by having the spin axis partly or completely perpendicular tothe solar wind flow during the deployment phase and keeping the voltageof a wire at a higher value when the wire is during its rotation movingalong the solar wind than when it Is moving in the opposite rotationphase back towards the solar wind flow. This principle is illustrated inFIG. 3, where the solar wind 301 now comes from the right. Wires shownas continuous lines, like wire 102, have their voltage turned higherthan wires shown as dashed lines, like wire 104. When part of the neededangular momentum is obtained from the solar wind, the auxiliarypropulsion system can be correspondingly smaller.

FIG. 4 summarizers schematically certain structural aspects of the wiredeployment and attachment technique. In the exemplary embodiment of FIG.4 the main body 101 of the spacecraft has initially a propulsionplatform 205 attached to it. The wire reels 201 are attached to thepropulsion platform 205 with an electrically isolating attachment 401.Attached to the main body 101 there is a wire holder 402, through whichthe wire 102 is deployed Into space. The attachment of the wire holder402 to the main body 101 is here shown to comprise an electricallyisolating mechanical attachment 403 and a controllable potentiometer404, which constitutes the only electrically conductive connectionbetween the wire 102 and the main body 101 of the spacecraft, and whichcan thus be used to individually control the voltage of each wire. Inthe lower portion of FIG. 4 the wire 102 has been fully deployed and isonly held by the wire holder 402. The auxiliary propulsion system andthe wire reels have been jettisoned.

Achieving and Controlling the Electric Potential of the Wires T

o maintain the wires at a positive potential, one must produce anelectric current from the surrounding plasma into the spacecraft.Emitting electrons with an electron gun is a straightforward andwell-known way of accomplishing this. The electricity needed by theelectron gun can be obtained from solar panels or some other onboardpower source. Using solar panels has the benefit that the electroncurrent that the wires and the main body of the spacecraft gather fromthe ambient solar wind plasma is proportional to the solar wind plasmadensity, which in turn is known to be, on the average, inverselyproportional to the squared distance from the Sun. The power producedwith solar panels scales in the same way so that the electric powerneeded versus the maximum power that can be produced by the panels isapproximately the same at all radial distances from the Sun. If someother power source having constant output power (like radiothermal powergenerator, RTG) is used, most of its output is not needed by thepropulsion system once the mission moves away from the Sun. Depending onthe mission, this may or may not be a benefit.

Electron current is gathered from the ambient plasma not only by thewires, but also by the spacecraft itself and Its structures (Includingsolar panels, if any). If a compact but otherwise arbitrarily shapedspacecraft of reasonable size would reside in space without the wires.It would create a nearly spherically symmetric potential pattern arounditself. In other words it would act approximately as a spherical probeembedded in plasma. Calculating the gathered current one finds that thespacecraft would gather so much electron current in this case thatkeeping it at large positive potential with an electron gun might bedifficult. When the radially extending, positively charged wires areadded to the system, however, the current gathered by the spacecraftdecreases dramatically because the potential pattern around thespacecraft and wires becomes disk-like. An electron approaching thesystem from outside is nearly as much attracted by the wire plane as bythe spacecraft body. When attracted by the wire plane, it is very likelyto move through the plane rather than hitting a wire because the wiresare very thin. Thus the radial, centrifugally supported wire plane thatis characteristic to this invention automatically solves the possibleproblem of prohibitively high gathered electron current.

Control and Navigation

FIGS. 5 a, 5 b and 5 c illustrate how the electric sail is used fornavigation. The spacecraft orbits the Sun on an elliptic orbit 501,according to the Keplerian laws of motion. The solar wind 301constitutes an essentially laminar flow in the direction perpendicularto the orbit in the orbit plane. The spacecraft spins around arotational axis 502 that here is also in the orbit plane. In FIG. 5 a weassume that the rotational axis 502 is initially perpendicular to theorbit, which means that the disklike electric sail faces directly thesolar wind 301. Neglecting the effect of the spinning motion at first,we also assume in FIG. 5 a that the spacecraft utilises thewire-specific potentiometers to turn on the voltage only in those wiresthat are on the right-hand half of the electric sail when looked at fromthe direction of the Sun. FIG. 5 a this is illustrated by representingthose wires with a continuous line that have their voltage turned on,and those wires with a dashed line that have their voltage turned off.The result is that the solar wind only exerts pressure on the right-handhalf of the electric sail, which in turn causes a torque that tends toturn the rotational axis 502 into the direction shown with a curvedarrow 503.

A force that tries to turn the rotational axis of a spinning bodyinteracts with the angular momentum of the body to cause a resultantthat acts in the direction that is perpendicular to both the originalrotational axis and the original force. This means that if the spin axisof the spinning electric-sail-propelled spacecraft is to be turned bychanging the potentials of the wires, the timing of the switching mustbe accomplished taking the angular momentum into account so that theeventual resultant force will point into the right direction. Anothermodification to the crude model considered above is that the wirepotentials should not be simply switched on and off, but to higher andlower values according to a suitable, carefully considered strategy.Using only “on” and “off” values would cause too large instantaneouschanges to the forces that act upon the wires.

A certain minimum amount of constant torque like that explained abovewould be needed anyway in most missions to keep the electric sail facingdirectly the solar wind, because without any torque the spinningspacecraft would tend to maintain the inertial orientation of itsrotational axis. However, we assume that the torque is larger than saidminimum value, so that the spacecraft will turn into the position shownIn FIG. 5 b In relation to its orbit. Now full voltage is applied to allwires. As a result, the dynamic pressure of the solar wind will cause aresultant force 51 1. This force has a component in the forward-lookingtangential direction of the orbit 501, which means that it increases theorbital velocity, causing the spacecraft to recede away from the Sun.The overall magnitude of the thrust Is proportional to the mean voltageof the wires, which in turn depends on the power used to operate theelectron gun.

Had the initial selection of the side on which the voltage is turned onbeen different in FIG. 5 a, the spacecraft would have turned into theposition shown in FIG. 5 b in relation to its orbit. Now the resultantforce 521 has a component in the backward-looking tangential directionof the orbit 501 which means that it decreases the orbital velocity,causing the spacecraft to assume a lower orbit closer to the Sun.

Since the voltage of each individual wire is separately controllable,oscillations in the wires that result from e.g. dynamic pressure pulsesin the solar wind and make the electric sail resemble a round sea raycan be damped by increasing and de creasing the wire voltages accordingto need. A control system on the spacecraft may detect such oscillationsthrough monitoring for example the impedances and tensions of the wiresor through mechanically monitoring the pointing direction of each wire,or through any other suitable means.

The large group of radially extending wires that constitute the electricsail is mechanically a very complex entity. Since the wires areextremely long compared to the dimensions of the spacecraft main body,the wire experiences the main body only as a point mass attached to oneend of the wire. After the initial moments of deployment the spinningmovement of the main body has an essentially negligible effect on theangular speed at which the wire rotates around the spinning axis.Inevitably some of the wires will have a tendency to speed up or slowdown in their angular motion, causing adjacent wires to touch each otherand get tangled.

One possible way of preventing the effect described above may be cleverenough control of wire potentials. The high electric potential of thesame sign in each wire will cause the wires to repel each other, andthis repulsion can be increased or decreased according to need bycontrolling the individual wire potentials. Another possibility is theso-called flight-time wire length control. This means using the wirereels or other available means to reel in or let out a length of theanomalously behaving wire. Reeling in some of the wire, i.e. decreasingthe length of the extended part of the wire, causes the remaining partto gain angular speed according to the principle of conserving angularmomentum. Correspondingly letting out an additional length of the wirewill slow down its angular speed. Introducing flight-time wire lengthcontrol naturally means that the wire reels cannot be jettisoned afterthe deployment phase.

The theoretical maximum value for the angle between radial outwarddirection and the direction of a wire extending into space from the rimof the spacecraft main body (see angle ψ in FIG. 2 b) is 90 degrees,because with a larger value the wire would start to wind around thespacecraft main body and/or touch the starting point of the adjacentwire. In practice the maximum allowable value must be somewhat less than90 degrees so that the wire does not get dangerously close to theadjacent wire at its starting point. It is advantageous to equip themain body with additional propulsive means capable of increasing anddecreasing the angular speed of the main body, so that by adding ortaking off some of the spinning motion of the main body, each radiallyextending wire of the electric sail can be kept pointing into adirection as close to directly radial as possible. Said additionalpropulsive means can be any known propulsion generators, including butnot being limited to chemical rockets, pressure release thrusters, ionengines and plasma engines.

System Level Considerations

FIG. 6 illustrates some of the subsystems of an exemplary spacecraftthat utilises an electric sail as propulsion means. The main controlresponsibility is on the mission control computer 601. A navigationcontrol system 602 controls the wire deploying mechanisms 603 and thewire voltage potentiometers 604 as well as the electron gun 605, thusdetermining the amount and direction of thrust that the spacecraft willget from the electric sail. Also the damping of oscillations and othertasks directly related to the wire voltages are on the responsibility ofthe navigation control system 602. It has a number of sensors 606 at itsdisposal, including for example accelerometers, sun sensors, forcedetectors, wire impedance detectors, wire pointing direction detectors,and an electron detector arranged to measure the potential of the mainbody of the spacecraft with respect to the surrounding plasma.

A spin rate control mechanism 610 is separately shown, with the purposeof providing the necessary Initial angular momentum for wire deploymentand increasing it during the wire deployment process. Shown also are theauxiliary propulsion mechanism 611, a spin rate measurement subsystem612 and the jettisoning mechanism 613, if any, that may be used to getrid of unnecessary mass after the wires have been fully deployed.

The power subsystem 621 uses solar panels and/or other energy sources toprovide the spacecraft, including the electron gun 605 with thenecessary operating power. A communications subsystem 622 enablescommunications with ground control. The spacecraft may have variouspayload subsystems 623.

A spacecraft that uses an electric sail as its primary propulsion systemin orbit could be used for various purposes, including Mercury-, Venus-or Sun-orbiting missions, fast fly by missions to outer solar systemsobjects and even interstellar missions outside the heliosphere, wherethe driving force is not any more solar wind but the flow ofinterstellar charged particles, the speed of which, however, is muchslower than that of solar wind within the heliosphere. An electric sailcould be used as braking means in a mission where the spacecraft wouldbe propelled out of the heliosphere with some other means, like laserassisted solar sailing.

An interesting possibility would be to place an electric-sail-propelledspacecraft to some point on the direct axis between the Earth and theSun. Previous missions of this kind have only involved placing aspacecraft to the so-called Lagrange point, in which the gravity of theEarth is just enough to keep the spacecraft on the proper orbit aroundthe Sun despite of its orbit velocity being otherwise too low. It takesonly about one hour for the solar wind to travel from the Lagrange pointto the Earth's magnetosphere, which means that a spacecraft measuringdisturbances in the solar wind there can not give a warning of changesin the solar weather very much in advance. An electric sail could enablea solar wind probe to hover at a location from which it takes e.g. fiveor six hours before the solar wind hits the Earth, which would give muchmore preparatory time for actions of protecting crews and equipment onspacecrafts orbiting the Earth. It would also increase dramatically theaccuracy of predicting activity in the aurora borealis phenomena.

The embodiments that we have described above are exemplary and do notlimit the applicability of the enclosed claims. For example, it is by nomeans mandatory to omit all transverse interconnections between wires,even if relying completely on radial wires will help to keep thedeployment procedure as simple as possible. The potentiometers or othercontrollable electric couplings between the electron gun and the wirescould be group-specific rather than wire-specific, so that the potentialof each wire group rather than each individual wire is controlled.s= Twoor more electron guns can be used as the electric potential generator onboard the spacecraft. The main body of the spacecraft does not need tobe disk-like, and the electric sail does not even need to be attached tothe main body; the electric sail may be located e.g. in a separatelyspinning subpart that is connected to a nonspinning main body. Onespacecraft may have two or more electric sails, much like somehelicopters have two main motors.

1. A spacecraft propulsion system, comprising: a plurality ofelectrically conductive elongated members adapted to be deployed from amain body into respective radial directions, an electric potentialgenerator adapted to generate an electric potential on board said mainbody, a controllable electric coupling between said electric potentialgenerator and said plurality of electrically conductive elongatedmembers, and an auxiliary propulsion system adapted to rotate said mainbody around a rotational axis that is perpendicular to said racialdirections.
 2. A spacecraft propulsion system according to claim 1,wherein each of said plurality of electrically conductive elongatedmembers Is one of the following: multifilament wire, braid, cable, band,tether.
 3. A spacecraft propulsion system according to claim 2,comprising a plurality of reels for storing said plurality ofelectrically conductive elongated members before deployment,
 4. Aspacecraft propulsion system according to claim 2, wherein there are notransverse couplings between adjacent ones of said plurality ofelectrically conductive elongated members.
 5. A spacecraft propulsionsystem according to claim 1, wherein said electric potential generatorcomprises an electron gun adapted to emit electrons from said main body.6. A spacecraft propulsion system according to claim 1, wherein saidcontrollable electric coupling comprises a separate controllableelectric coupling between said main body and each of said plurality ofelectrically conductive elongated members.
 7. A spacecraft propulsionsystem according to claim 6, wherein said controllable electriccouplings are electrically controllable potentiometers.
 8. A spacecraftpropulsion system according to claim 1, comprising propulsive unitsadapted to controllably produce an angular momentum around saidrotational axis.
 9. A spacecraft propulsion system according to claim 8,wherein said propulsive units are located in a detachable platform. 10.A spacecraft propulsion system according to claim 9, comprising storagemechanisms for storing said plurality of electrically conductiveelongated members before deployment, wherein said storage mechanisms arelocated in said detachable platform.
 11. A spacecraft propulsion systemaccording to claim 1, comprising a navigation system adapted to controlsaid controllable electric coupling in order to change the attitude ofthe spacecraft in relation to a surrounding flow of charged particles.12. A spacecraft propulsion system according to claim 1, comprising anavigation system adapted to controllably change the extended length ofindividual electrically conductive elongated members.
 13. A spacecraftpropulsion system according to claim 1, comprising sensors adapted tosense mechanical oscillations in said plurality of electricallyconductive elongated members, and a control system adapted to controlsaid controllable electric coupling in order to damp detected mechanicaloscillations.
 14. A method for changing the state of motion of aspacecraft from Its natural Keplerian motion, comprising: generating anelectric potential on board a main body of the spacecraft, controllingan electric coupling between the generated electric potential and aplurality of electrically conductive elongated members deployed fromsaid main body into respective radial directions, and rotating said mainbody around a rotational axis that is perpendicular to said radialdirections in order to cause a centrifugal tensile force to said aplurality of electrically conductive elongated members.
 15. A methodaccording to claim 14, wherein a subset of said electrically conductiveelongated members is dynamically kept at a positive potential withrespect to a flow of charged particles around the spacecraft, in orderto create a torque that modifies a rotation state of the spacecraft. 16.A method according to claim 15, wherein: said rotational axis is keptperpendicular to said flow of charged particles around the spacecraft,and said subset of said electrically conductive elongated membersconsists of such electrically conductive elongated members that due tothe spacecraft's rotation around said rotational axis move along withsaid flow of charged particles around the spacecraft, in order toincrease the spacecraft's angular momentum around said rotational axis.17. A method according to claim 16, wherein the method steps of claim 16are performed during a deployment phase of said electrically conductiveelongated members from the main body of the spacecraft, and thecentrifugal force associated with said angular momentum assists in thedeployment.
 18. A method according to claim 15, wherein: said rotationalaxis Is kept in an orbital plane of motion of the spacecraft andnonperpendicular to said flow of charged particles around thespacecraft, said subset of said electrically conductive elongatedmembers consists of an essentially equal number of electricallyconductive elongated members above and below said orbital plane ofmotion, and all electrically conductive elongated members of said subsetare on the same side, which is either the leading side or the trailingside, of the main body of the spacecraft in relation to its orbitalmotion, in order to turn said rotational axis in said orbital plane ofmotion.
 19. A method according to claim 14, comprising controllablychanging the ex-tended length of individual electrically conductiveelongated members after an initial deployment of said electricallyconductive elongated members, in order to controllably change angularspeeds of individual electrically conductive elongated members aroundsaid rotational axis.
 20. A computer program product on acomputer-readable medium, comprising software instructions that, whenexecuted in a computer system, cause the implementation of: controllingthe generation of an electric potential on board a main body of aspacecraft, controlling an electric coupling between the generatedelectric potential and a plurality of electrically conductive elongatedmembers deployed from said main body into respective radial directions,and controlling the rotation of said main body around a rotational axisthat is perpendicular to said radial directions in order to cause atensile force to said plurality of electrically conductive elongatedmembers through centrifugal motion.
 21. A computer program productaccording to claim 20, comprising software instructions that, whenexecuted in a computer system, cause the implementation of controllablychanging the extended length of individual electrically conductiveelongated members after an initial deployment of said electricallyconductive elongated members, in order to controllably change angularspeeds of individual electrically conductive elongated members aroundsaid rotational axis.